Porous ceramic coating system

ABSTRACT

A coating system having two layers of porous coatings is provided. The layers include a tightly controlled and matched porosity. The layer system also includes a substrate with a bonding layer. The inner ceramic layer and the outer ceramic layer are formed on the bonding layer. The bonding layer may comprise a MCrAlX alloy.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is the US National Stage of International ApplicationNo. PCT/EP2011/061320, filed Jul. 5, 2011 and claims the benefitthereof. The International Application claims the benefits of EuropeanPatent Office application No. 10007240.4 EP filed Jul. 14, 2010. All ofthe applications are incorporated by reference herein in their entirety.

FIELD OF INVENTION

The invention relates to a layer system having two different porousceramic layers.

BACKGROUND OF INVENTION

Ceramic protective layers are often used for components used at hightemperatures in order to protect the metallic substrate from relativelyhigh temperatures.

In this respect, the ceramic layers have a certain porosity in orderfirstly to reduce the thermal conductivity and in order to set a certainductility.

SUMMARY OF INVENTION

It is an object of the invention to optimize the thermal and themechanical properties.

The object is achieved by a layer system as claimed in the claims

The dependent claims list further advantageous measures which can becombined with one another, as desired.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a layer system,

FIG. 2 shows a gas turbine,

FIG. 3 shows a turbine blade or vane,

FIG. 4 shows a combustion chamber,

FIG. 5 shows a list of superalloys.

The description and the figures show merely exemplary embodiments of theinvention.

DETAILED DESCRIPTION OF INVENTION

FIG. 1 schematically shows the layer system. The layer system 1 ispreferably a turbine blade or vane 120, 130 of a turbine, of a steamturbine, of a gas turbine 100 (FIG. 2), for stationary operation or foraircraft.

It is preferable for the substrate 4 to comprise a nickel-based orcobalt-based superalloy made of an alloy shown in FIG. 5. This ispreferably a nickel-based superalloy.

A metallic bonding layer, in particular of the MCrAl or MCrAlX type(M═Ni, Co, Fe, preferably Ni, Co), is preferably present on thesubstrate 4.

Similarly, there may be diffusion layers within the substrate 4, towhich a ceramic layer 16 can be applied.

A ceramic layer 16 is applied to the metallic layer 7 or to thesubstrate 4, wherein an oxide layer (TGO) is either produceddeliberately on or applied to the interface, this forming during theceramic coating or during operation of a layer system having themetallic layer 7.

The ceramic layer 16 has at least two, in particular only two, differentceramic layers 10, 13. The lower ceramic layer 10 has a lower porositythan the outer ceramic layer 13. The porosity of the lower ceramic layer10 is 8% to 14%, in particular 11% to 12% (preferably % by volume).

It is preferable for the layer thickness of the inner ceramic layer 10to be at least 10%, in particular 20%, very particularly 50%, thinnerthan that of the outer ceramic layer 13. The lower layer has a thicknessof 100±25 μm, whereas the outer layer has a thickness of >100 μm.

The outer ceramic layer 13 has a porosity of 14% to 18% and ispreferably the outermost layer, which is exposed directly to the hotgas.

The material of the lower ceramic layer 10 is partially stabilized, inparticular by yttrium, zirconium oxide. This material is preferably alsoused for the outer ceramic layer 13, although it is also possible to usea pyrochlore material.

The selection of the porosity of the ceramic layers surprisingly led toa longer service life compared to a layer of high porosity and equalthickness.

FIG. 2 shows, by way of example, a partial longitudinal section througha gas turbine 100.

In the interior, the gas turbine 100 has a rotor 103 with a shaft whichis mounted such that it can rotate about an axis of rotation 102 and isalso referred to as the turbine rotor.

An intake housing 104, a compressor 105, a, for example, toroidalcombustion chamber 110, in particular an annular combustion chamber,with a plurality of coaxially arranged burners 107, a turbine 108 andthe exhaust-gas housing 109 follow one another along the rotor 103.

The annular combustion chamber 110 is in communication with a, forexample, annular hot-gas passage 111, where, by way of example, foursuccessive turbine stages 112 form the turbine 108.

Each turbine stage 112 is formed, for example, from two blade or vanerings. As seen in the direction of flow of a working medium 113, in thehot-gas passage 111 a row of guide vanes 115 is followed by a row 125formed from rotor blades 120.

The guide vanes 130 are secured to an inner housing 138 of a stator 143,whereas the rotor blades 120 of a row 125 are fitted to the rotor 103for example by means of a turbine disk 133.

A generator (not shown) is coupled to the rotor 103.

While the gas turbine 100 is operating, the compressor 105 sucks in air135 through the intake housing 104 and compresses it. The compressed airprovided at the turbine-side end of the compressor 105 is passed to theburners 107, where it is mixed with a fuel. The mix is then burnt in thecombustion chamber 110, forming the working medium 113. From there, theworking medium 113 flows along the hot-gas passage 111 past the guidevanes 130 and the rotor blades 120. The working medium 113 is expandedat the rotor blades 120, transferring its momentum, so that the rotorblades 120 drive the rotor 103 and the latter in turn drives thegenerator coupled to it.

While the gas turbine 100 is operating, the components which are exposedto the hot working medium 113 are subject to thermal stresses. The guidevanes 130 and rotor blades 120 of the first turbine stage 112, as seenin the direction of flow of the working medium 113, together with theheat shield elements which line the annular combustion chamber 110, aresubject to the highest thermal stresses.

To be able to withstand the temperatures which prevail there, they maybe cooled by means of a coolant.

Substrates of the components may likewise have a directional structure,i.e. they are in single-crystal form (SX structure) or have onlylongitudinally oriented grains (DS structure).

By way of example, iron-based, nickel-based or cobalt-based superalloysare used as material for the components, in particular for the turbineblade or vane 120, 130 and components of the combustion chamber 110.

Superalloys of this type are known, for example, from EP 1 204 776 B1,EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949.

The guide vane 130 has a guide vane root (not shown here), which facesthe inner housing 138 of the turbine 108, and a guide vane head which isat the opposite end from the guide vane root. The guide vane head facesthe rotor 103 and is fixed to a securing ring 140 of the stator 143.

FIG. 3 shows a perspective view of a rotor blade 120 or guide vane 130of a turbomachine, which extends along a longitudinal axis 121.

The turbomachine may be a gas turbine of an aircraft or of a power plantfor generating electricity, a steam turbine or a compressor.

The blade or vane 120, 130 has, in succession along the longitudinalaxis 121, a securing region 400, an adjoining blade or vane platform 403and a main blade or vane part 406 and a blade or vane tip 415.

As a guide vane 130, the vane 130 may have a further platform (notshown) at its vane tip 415.

A blade or vane root 183, which is used to secure the rotor blades 120,130 to a shaft or a disk (not shown), is formed in the securing region400.

The blade or vane root 183 is designed, for example, in hammerhead form.Other configurations, such as a fir-tree or dovetail root, are possible.

The blade or vane 120, 130 has a leading edge 409 and a trailing edge412 for a medium which flows past the main blade or vane part 406.

In the case of conventional blades or vanes 120, 130, by way of examplesolid metallic materials, in particular superalloys, are used in allregions 400, 403, 406 of the blade or vane 120, 130.

Superalloys of this type are known, for example, from EP 1 204 776 B1,EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949.

The blade or vane 120, 130 may in this case be produced by a castingprocess, by means of directional solidification, by a forging process,by a milling process or combinations thereof.

Workpieces with a single-crystal structure or structures are used ascomponents for machines which, in operation, are exposed to highmechanical, thermal and/or chemical stresses.

Single-crystal workpieces of this type are produced, for example, bydirectional solidification from the melt. This involves castingprocesses in which the liquid metallic alloy solidifies to form thesingle-crystal structure, i.e. the single-crystal workpiece, orsolidifies directionally.

In this case, dendritic crystals are oriented along the direction ofheat flow and form either a columnar crystalline grain structure (i.e.grains which run over the entire length of the workpiece and arereferred to here, in accordance with the language customarily used, asdirectionally solidified) or a single-crystal structure, i.e. the entireworkpiece consists of one single crystal. In these processes, atransition to globular (polycrystalline) solidification needs to beavoided, since non-directional growth inevitably forms transverse andlongitudinal grain boundaries, which negate the favorable properties ofthe directionally solidified or single-crystal component.

Where the text refers in general terms to directionally solidifiedmicrostructures, this is to be understood as meaning both singlecrystals, which do not have any grain boundaries or at most havesmall-angle grain boundaries, and columnar crystal structures, which dohave grain boundaries running in the longitudinal direction but do nothave any transverse grain boundaries. This second form of crystallinestructures is also described as directionally solidified microstructures(directionally solidified structures).

Processes of this type are known from U.S. Pat. No. 6,024,792 and EP 0892 090 A1.

The blades or vanes 120, 130 may likewise have coatings protectingagainst corrosion or oxidation e.g. (MCrAlX; M is at least one elementselected from the group consisting of iron (Fe), cobalt (Co), nickel(Ni), X is an active element and stands for yttrium (Y) and/or siliconand/or at least one rare earth element, or hafnium (Hf)). Alloys of thistype are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 orEP 1 306 454 A1.

The density is preferably 95% of the theoretical density.

A protective aluminum oxide layer (TGO=thermally grown oxide layer) isformed on the MCrAlX layer (as an intermediate layer or as the outermostlayer).

The layer preferably has a composition Co-30Ni-28Cr-8Al-0.6Y-0.7Si orCo-28Ni-24Cr-10Al-0.6Y. In addition to these cobalt-based protectivecoatings, it is also preferable to use nickel-based protective layers,such as Ni-10Cr-12Al-0.6Y-3Re or Ni-12Co-21Cr-11Al-0.4Y-2Re orNi-25Co-17Cr-10Al-0.4Y-1.5Re.

It is also possible for a thermal barrier coating, which is preferablythe outermost layer, to be present on the MCrAlX, consisting for exampleof ZrO₂, Y₂O₃-ZrO₂, i.e. unstabilized, partially stabilized or fullystabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide.

The thermal barrier coating covers the entire MCrAlX layer. Columnargrains are produced in the thermal barrier coating by suitable coatingprocesses, such as for example electron beam physical vapor deposition(EB-PVD).

Other coating processes are possible, e.g. atmospheric plasma spraying(APS), LPPS, VPS or CVD. The thermal barrier coating may include grainsthat are porous or have micro-cracks or macro-cracks, in order toimprove the resistance to thermal shocks. The thermal barrier coating istherefore preferably more porous than the MCrAlX layer.

The blade or vane 120, 130 may be hollow or solid in form.

If the blade or vane 120, 130 is to be cooled, it is hollow and may alsohave film-cooling holes 418 (indicated by dashed lines).

FIG. 4 shows a combustion chamber 110 of the gas turbine 100.

The combustion chamber 110 is configured, for example, as what is knownas an annular combustion chamber, in which a multiplicity of burners107, which generate flames 156, arranged circumferentially around anaxis of rotation 102 open out into a common combustion chamber space154. For this purpose, the combustion chamber 110 overall is of annularconfiguration positioned around the axis of rotation 102.

To achieve a relatively high efficiency, the combustion chamber 110 isdesigned for a relatively high temperature of the working medium M ofapproximately 1000° C. to 1600° C. To allow a relatively long servicelife even with these operating parameters, which are unfavorable for thematerials, the combustion chamber wall 153 is provided, on its sidewhich faces the working medium M, with an inner lining formed from heatshield elements 155.

Moreover, a cooling system may be provided for the heat shield elements155 and/or their holding elements, on account of the high temperaturesin the interior of the combustion chamber 110. The heat shield elements155 are then, for example, hollow and may also have cooling holes (notshown) opening out into the combustion chamber space 154.

On the working medium side, each heat shield element 155 made from analloy is equipped with a particularly heat-resistant protective layer(MCrAlX layer and/or ceramic coating) or is made from material that isable to withstand high temperatures (solid ceramic bricks).

These protective layers may be similar to the turbine blades or vanes,i.e. for example MCrAlX: M is at least one element selected from thegroup consisting of iron (Fe), cobalt (Co), nickel (Ni), X is an activeelement and stands for yttrium (Y) and/or silicon and/or at least onerare earth element or hafnium (Hf). Alloys of this type are known fromEP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.

It is also possible for a, for example ceramic, thermal barrier coatingto be present on the MCrAlX, consisting for example of ZrO₂, Y₂O₃—ZrO₂,i.e. unstabilized, partially stabilized or fully stabilized by yttriumoxide and/or calcium oxide and/or magnesium oxide.

Columnar grains are produced in the thermal barrier coating by suitablecoating processes, such as for example electron beam physical vapordeposition (EB-PVD).

Other coating processes are possible, e.g. atmospheric plasma spraying(APS), LPPS, VPS or CVD. The thermal barrier coating may include grainsthat are porous or have micro-cracks or macro-cracks, in order toimprove the resistance to thermal shocks.

Refurbishment means that after they have been used, protective layersmay have to be removed from turbine blades or vanes 120, 130 or heatshield elements 155 (e.g. by sand-blasting). Then, the corrosion and/oroxidation layers and products are removed. If appropriate, cracks in theturbine blade or vane 120, 130 or the heat shield element 155 are alsorepaired. This is followed by recoating of the turbine blades or vanes120, 130 or heat shield elements 155, after which the turbine blades orvanes 120, 130 or the heat shield elements 155 can be reused.

1-8. (canceled)
 9. A layer system, comprising: a substrate; a metallicbonding layer on the substrate; an oxide layer on the bonding layer oron the substrate; an inner ceramic layer having a porosity of 11% to12%; and an outermost ceramic layer, having a porosity of 16% to 18%, onthe inner ceramic layer.
 10. The layer system as claimed in claim 9,wherein the metallic bonding layer comprises an MCrAl or MCrAlX alloy.11. The layer system as claimed in claim 9, wherein the metallic bondinglayer consists of an MCrAl or MCrAlX alloy.
 12. The layer system asclaimed in claim 9, wherein the material of the inner ceramic layercomprises zirconium oxide.
 13. The layer system as claimed in claim 9,wherein the material of the inner ceramic layer comprises partiallystabilized zirconium oxide.
 14. The layer system as claimed in claim 9,wherein the material of the inner ceramic layer comprises zirconiumoxide partially stabilized by yttrium.
 15. The layer system as claimedin claim 9, wherein the material of the inner ceramic layer consists ofzirconium oxide partially stabilized by yttrium.
 16. The layer system asclaimed in claim 9, wherein the outer layer comprises zirconium oxide.17. The layer system as claimed in claim 9, wherein the outer layercomprises partially stabilized zirconium oxide.
 18. The layer system asclaimed in claim 9, wherein the outer layer comprises zirconium oxidepartially stabilized by yttrium.
 19. The layer system as claimed inclaim 9, wherein the outer layer consists of zirconium oxide partiallystabilized by yttrium.
 20. The layer system as claimed in claim 9,wherein the substrate comprises a nickel-based or cobalt-basedsuperalloy.
 21. The layer system as claimed in claim 9, wherein thesubstrate consists of a nickel-based or cobalt-based superalloy.
 22. Thelayer system as claimed in claim 9, wherein the material of the innerceramic layer and the material of the outer ceramic layer are different.23. The layer system as claimed in claim 22, wherein the outer ceramiclayer comprises a pyrochlore structure.
 24. The layer system as claimedin claim 9, wherein the inner ceramic layer is at least 10% thinner thanthe outermost ceramic layer.
 25. A layer system, consisting of: asubstrate; a metallic bonding layer; an oxide layer on the bondinglayer; an inner ceramic layer; and an outermost ceramic layer.